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Figure 2


Fig. 2. Lift and drag coefficients. (A) An airfoil (NACA2414) at Re~3x105. Circles are experimental data (Selig, 2002) and solid lines are given by CL=2.77sin2({alpha}+0.03), CD=0.0086+0.24[1–cos2({alpha}–0.02)]. (B) A low Reynolds number plate at Re~103, CL=1.5sin2{alpha}, CD=1.1–cos2{alpha}.





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